Gas Turbine Characteristics for a Large Civil Tilt-Rotor ...

NASA/TM-2010-216089

Gas Turbine Characteristics for aLarge Civil Tilt-Rotor (LCTR)

Christopher A. SnyderGlenn Research Center, Cleveland, Ohio

Douglas R. Thin-manU.S. Army Research Laboratory, Glenn Research Center; Cleveland, Ohio

February 2010

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NASA/TM-2010-216089

Gas Turbine Characteristics for aLarge Civil Tilt-Rotor (LCTR)

Christopher A. SnyderGlenn Research Center, Cleveland, Ohio

Douglas R. Thin-manU.S. Army Research Laboratory, Glenn Research Center; Cleveland, Ohio

Prepared for the65th Annual Forum and Technology Display (AHS Forum 65)sponsored by the American Helicopter SocietyGrapevine, Texas, May 27-29, 2009

National Aeronautics andSpace Administration

Glenn Research CenterCleveland, Ohio 44135

February 2010

This report is a formal draft or workingpaper, intended to solicit comments and

ideas from a technical peer group.

This report contains preliminary findings,subject to revision as analysis proceeds.

Trade names and trademarks are used in this report for identificationonly. Their usage does not constitute an official endorsement,either expressed or implied, by the National Aeronautics and

Space Administration.

Level of Review: This material has been technically reviewed by technical management

Available from

NASA Center for Aerospace Information National Technical Information Service7115 Standard Drive 5285 Port Royal RoadHanover, MD 21076-1320 Springfield, VA 22161

Available electronically at http://gltrs.grc.nasa.gov

Gas Turbine Characteristics for a Large Civil Tilt-Rotor (LCTR)

Christopher A. SnyderNational Aeronautics and Space Administration


Douglas R. ThunmanU.S. Army Research Laboratory


AbstractIn support of the Fundamental Aeronautics Program,

Subsonic Rotary Wing Project, an engine system study hasbeen undertaken to help define and understand some of themajor gas turbine engine parameters required to meetperformance and weight requirements as defined by earliervehicle system studies. These previous vehicle studies will bereviewed to help define gas turbine performance goals.Assumptions and analysis methods used will be described.Performance and wei ght estimates for a few conceptual gasturbine engines meeting these requirements will be given anddiscussed. Estimated performance for these conceptualengines over a wide speed variation (down to 50 percentpower turbine rpm at hi gh torque) will be presented. Finally,areas needing further effort will be suggested and discussed.

Nomenclature

C.G. center of gravity, incheseff efficiencyfps feet per secondft feethp horsepowerHPC high-pressure compressorHPT high-pressure turbinelbm pounds massLCTR Large Civil Tilt RotorLP low pressureLPC low-pressure compressorLPT low-pressure turbineMax maximumN actual speed, rpmNc corrected speed, NIJO , rpmPR pressure ratioPSFC power specific fuel consumption, lbm/hr/lipPT power turbine

sec secondT3 compression system exit temperature, °FT4 combustor exit temperature, °FVtip rotor tip velocity, feet per secondW actual mass flow, lbm/secWe corrected mass flow, W * FO/8 , lbm/secWturb power turbine actual mass flow, lbm/sec

ratio of actual to standard pressureratio of actual to standard temperature

IntroductionThe NASA Heavy Lift Rotorcraft System Investigation

(Ref. 1) identified a large tilt rotor as the best concept to meetthe various airspace and other requirements for the future,short-haul regional market. This evolved into a conceptualvehicle designated as LCTR2 (Large Civil Tilt Rotor—iteration 2) (Ref. 2) as seen in Figure 1

This vehicle iteration was designed to carry 90 passengersat 300 knots with at least a 1,000 n nu range; powered by fourturboshaft engines designed for 7,500 hp each. Other designfeatures included a rotor tip speed of 650 ft/sec in hover and350 ft/sec durin g cruise, enabled by a two-speed gearbox. Thisrange of rotor tip speeds was needed to achieve the high levelof performance and efficiency at two very different flightconditions. The rotor tip speed variation could theoretically beobtained usin g a variable diameter rotor or multiple-speedgearboxes (orVa combination of these or other approaches);this work is focusing on achieving all speed variation from theengine. Although the exact requirements and characteristicsfor such a vehicle class are still being researched, performinganalyses on a representative vehicle will help understand thesensitivities for such a design, help guide research efforts toreduce risks, and develop a suite of technologies from whichthis new vehicle class and capability can evolve and bedeveloped. The final vehicle design could use one or acombination of these variable rotor tip speed concepts,determined from the vehicle's specific design and nussionrequirements and the state of these various requiredtechnologies.

NASA/TM-2010-216089

Figure 1.—Conceptual view of LCTR2.

This report details gas turbine engine technologyassumptions and analyses performed to estimate the engineparameters needed to obtain sufficient performance to meetoperational goals for the proposed vehicle concept. Thisvehicle would use the turboshaft version of a gas turbineengine. The core gas turbine engine develops high energy(high temperature and pressure) gas to power a separate powerturbine and shaft that supplies horsepower and torque to themain drive system. Initial cycle parametrics were performed tosuggest gas turbine engine characteristics that would meetLCTR2 performance requirements. Based on these overallcharacteristics, therniodynamic engine analyses assumin g aone-spool core (all compression on one shaft) and a two-spoolcore (with the compression split between two sets ofcompressors, each on their own shaft) were performed. Asstated above, both configurations have a separate powerturbine and shaft. For the thermodynamic analyses,compressor component performance maps were generated andused to estimate off-design performance at the hover andcruise flight points. At the cniise condition, gas turbine engineperformance was estimated at 100, 75 and 50 percent powerturbine rpm to quantify engine performance at reduced powerturbine speeds. Results will be discussed and suggestions forfarther analysis will be given. Follow-on studies are underwayto perform more detailed analyses of compression and turbinesystems and component performance. It is expected that themore detailed component analyses will be incorporated andreported in subsequent studies and reports. It must be notedthat this work is not expected to identify the specific gasturbine engine attributes and cycle definitions to meet allrequirements. Its purpose is to further refine requirements andidentify possible components, systems or subsystems thatcould enable such new classes of engine and vehicle designsand operations while also uncovering areas requiring furtherexploration and development.

Analysis MethodologyEngine system studies were performed to estimate the

major gas turbine engine parameters that would meetperformance requirements for the previously defined vehicleand mission. Engine power, weight and fuel consumption are

important performance parameters that will help definepossible engine configurations. As part of the parametricanalysis, compressor pressure ratio (PR) was varied from 5 to60, assuming a constant polytropic efficiency of 88 percent.Although turbomachinery efficiency would vary dependin g onengine size and configuration, that effect was deferred to lateranalyses. Combustor exit temperature was varied from 2000 to3200 °F (in 400° increments). To get turbine cooling bleedestimates for the core turbines, the method of Gauntner

(Ref. 3) was used, assuming metal temperatures of 2200 °Ffor the stator, 2100 °F for the rotor. These turbine metaltemperatures are higher than present, small gas turbines toinclude the effects of incorporation of improved turbinematerial temperature capabilities and cooling techniques inthese smaller turbine sizes. The power turbine was assumed tobe urncooled. These should be reasonable temperatures andefficiencies for engines in the LCTR2-size class with entry-in-service in roughly the 2020 timeframe.

The object-oriented analysis framework, the Numerical

Propulsion System Simulator (NPSS) (Ref. 4), was used toperform the gas turbine analyses. NPSS contains standard 0/1-D elements for the gas turbine components. These areconfigured into a representative steady-state, thermodynamicmodel. An example block dia gram representative of a one-spool core, turboshaft is shown in Figure 2. Further elementsare defined to drive specific parameters to desired values andinsure continuity of mass, momentum and ener gy. After initialcycle parameters were determined, CMGEN (Ref. 5) was usedto generate compressor performance maps that would be morerepresentative of the flog-speed characteristics for a givencompressor PR and size during the off-design analyses.

The gas turbine flow path and weight were generated usingthe WATE (Ref. 6) program. Using the output from NPSS(mass flows, temperatures, pressures, velocities, etc.) andfurther user input, A"ATE sizes the various mechanical andflow components for the gas turbine engine, determiningmaterials, dimensions and weights for the differentcomponents represented. As part of the process, WATE alsoproduces a graphical representation that can be used to checkfor reasonable component dimensions and ensure that thereare no discontinuities or sharp turns in the gas flow path. Theresults of these analyses also form the basis for more detailedfollow-on studies.

Results and Discussion

Initial Engine Parametrics

The preliminary analysis is instructive to suggest engineparameters to meet vehicle and mission requirements. Figure 3shows calculated Power Specific Fuel Consumption (PSFC-lbm/hr fuel per hp produced) versus high pressure compressor(HPC) PR and combustor exit temperature (T4). Also includedin the graph are areas representative of the performanceregions for the Honeywell Aerospace T5.5 and GE Aviation

NASA/TM-2010-216089

HPC bleeds for turbine cooli

High-pressure High-

PowerInlet Duct Duct com- Diffuser Combustor pressure Duct Nozzle

pressor I. turbine turbine

High-pressure shaftLow-pressure shaft

T3 T4 Horsepower Drive ♦ Mainextraction system rotor

Figure 2.—Block representation of a one-spool core turboshaft gas turbine model.

0.600

CL0.550

E0.500

00.450

0

0.400U_

0.3500 10 20 30 40 50 60

HPC pressure ratio

Figure 3.—Power Specific Fuel Consumption (PSFC, Ibm/hr/hp) versus compressorpressure ratio and combustor exit temperature, T4.

T700 families of turboshaft engines, the version of RollsRoyce AE 1107 in the Bell Boeing V-22 aircraft (all fromRef 7) and the estimated level of PSFC used in the LCTR2system studies. With the assumed technology levels, there is aminimum in the PSFC curves at a HPC PR of 30 (for values ofT4 around 2800 to 3200 °F—both curves fall almost on top ofeach other). To meet the LCTR2 design of 7,500 hp per enginein that region of T4 and HPC PR would require an airflow ofapproximately 30 lbm/sec. Therefore, based on the technologyassumptions, the engine parameters (at the sea level staticdesign) used for subsequent analyses were: airflow of30 lbrri/sec, overall pressure ratio of 30, and T4 of 3000 °F.Table 1 compares the major parameters for this gas turbine tothe Rolls Royce AE1107 (the engine in the Bell Boeing V-22rotorcraft, a modern turboshaft engine in a slightly smallerpower class). To meet the PSFC requirements ; the notionalstudy engine will need to operate at temperatures andpressures significantly higher than those found in presentturboshaft engines ; but at levels already in modern, largeengines ; while maintaining compressor and turbine efficiencyand performance at the required much smaller airflows andblade sizes. A few conceptual engine configurations wereinvestigated and will now be discussed in concert with theirrepresentative engine thermodynamic and flow patharrangements.

TABLE 1.—GAS TURBINE ENGINE PARAMETERSParameter AE1107 Notional engineHorsepower 6,000 7,500,Weight, Ibm 971 1,000,Airflow, lbm/sec 353 30PSFC, lbun/hr/hp .426 .3 7'Overall pressure ratio 16.7 30Compressor exit temperature, T3, °F 810 1099Combustor exit temperature, T4, T 2200 3000Corrected flow:

Compressor entrance 35.5 30Compressor exit 3.2 1.4

'Parameters are from Reference 2.

One-Spool Core Engine

The first engine to be modeled was a one-spool core (with afree turbine on a second spool) with an all-axial compressor.This allowed the engine flow path to be laid out and togenerate compressor component perfonnance maps to do aninitial check on off-design performance. Compressor mapswere generated at the design pressure ratio and flow.Two nussion profiles were evaluated, each with a hovercondition and a range of conditions at cruise. Gas turbineperformance was calculated at the key points (hover and initialcruise power), to verify sufficient engine horsepowerwas available at full power turbine rpm (hover) and at

NASA/TM-2010-216089

100 percent, 7.5 percent and the reduced power turbinemechanical rpm identified for optimum rotor performance atcruise. Results are given in Table 2 detailing the flightcondition and pertinent turbomachinery operatingcharacteristics. This preliminary analysis indicated that the gasturbine was able to achieve horsepower requirements for allflight points except one. There was insufficient enginehorsepower for the mission 2 initial cruise point (operating athigher than the 300 knot requirement at reduced rotor / power

turbine rpm). However, the engine was able to meet vehiclepower requirements after about 1/3 of cruise fuel wasconsumed, reducing the vehicle weight and power needed. Itwas also found that at the cruise condition, 100 percent powerturbine mechanical speed actually results in an aerodynamicover speed as a result of the drop in ambient temperature withaltitude. This suggests that even if a multispeed gearbox wasused, power turbine rpm could need to be reduced for cruiseoperation.

TABLE 2.-OPERATION OF ONE-SPOOL CORE TURBOSHAFT ENGINE OVERREFERENCE MISSION PROFILES AT VARIOUS POWER TURBINE RPMS

Altitude, Speed. Vtip, fps AT. Meet Core rpm HPC HPT Ne T4, °F PT Ne PIP/ We-PTft knots (rotor) °F ]]p`' (actual,/design) Nc Wtrrb° (lbnl/s)0 0 650 +27 Yes 1.02 99.2 1 100.4 3091 98.5 1 261.7 12.54

Mission 12.000 0 650 -45 Yes 0.99 95.6 101.0 2864 101.8 215.9 12.16

(100% PT rpm) 28,000 303.4 350 0 Yes 0.90 98.0 99.9 2367 111.7' 210.2 12.60(75% PT rpm) Yes 0.90 98.1 100.3 2352 84.1'' 208.4 12.86(50% PT rpm) Yes 0.91 99.3 100.5 2405 59.9 198.9 13.21

Mission 25,000 0 650 -36 Yes 0.99 97.0 100.6 2870 101.9 227.1 12.31

(100% PT rprn) 28,000 330 350 0 Yes 0.92 99.1 99.7 2464 109.7" 228.7 12.71(75% PT rpni) Yes 0.92 99.4 100.2 2459 82.5 225.5 13.00(50% PT spin) NO 0.93 100.0 100.4 3480 59.1 308.8 13.26

(50% PT rpm) End 28,000 330 350 0 Yes 0.92 99.9 100.4 2472 59.2 207.9 13.26'Actual power turbine horsepower /mass flow (hp/lbm/sec)bMultispeed gearbox required to match rotor and power turbine speeds)

The operating points for the core turbomachinery andpower turbine are shown in Figures 4 to 6. As can be seen inFigure 4 and Figure 5, the core compressor and turbine operatein a fairly narrow band over the fli ght profile. As shown inFigure 6, the power turbine has a much larger variation in itsrpm. Although power turbine corrected flow varies less than 6percent from its sea level design value ; its pressure ratioincreases about 1.5 percent from the full speed rpm at hover tothe part speed rpm of the cruise. As shown in Table 2, theamount of engine power required per pound of actual powerturbine mass flow is fairly constant for both missions at bothflight points. This was not unexpected, since flight altitude andspeed are determined by the engine power available (amongother factors), with actual engine airflow and power fallingwith the increase in altitude and the accompanying drop inambient density_ However ; it was hoped that powerrequirements per lbm airflow would drop to allow powerturbine rpm to decrease as the mission progressed from hoverto cruise, such that the main rotor rpm could also be reducedto maintain high rotor efficiency. For this analysis, at reducedrotor and power turbine rpm, the loss of turbine efficiency iscompensated by an increase in pressure ratio to maintainpower production. Maintaining a constant power turbinehorsepower output per lbm airflow with an almost 50 percentreduction in rpm without loss of efficiency requires turbinedesign unique to this vehicle and nussion class. A preliminaryanalysis looking at variable vanes and blades to meet this

requirement is given in Reference 8. One such design conceptincludes power turbine variable incident nozzles, incidenttolerant blades and additional stages. Whatever the designsolution, further analysis is needed for the power turbine toverify its performance at almost constant power to weight flowover such a range of rotational speeds while still meetingengine weight goals.

Engine gas flow path dimensions and weights weregenerated assuming an all axial compression system, outputfrom the WATE analysis is shown in Figure 7. It resulted in anengine plus accessories total weight of 946 lb (about 5 percentlighter than assumed in the LCTR2 study). At this engineairflow and horsepower class with one spool, all axialcompression results in 10 stages to achieve the desiredcompression (tip speeds are below 1,000 ft per second for thelast 4 stages). This specific application also results in verysmall blade heights for the latter stages (less than 1 in. for thelast 4 stages, the last stage blade height is only 0.57 in.). Thissize blade could be a challenge for achieving and maintainingefficiency over the life of the engine. Further discussion onthis particular design challenge is given in Reference 9. Thepower turbine WATE results assume typical turbine stagedesign. If additional power turbine stages using incidenttolerant blades (and other variability) were needed to maintaingood operation over its large speed range, these factors couldincrease power turbine weight.

NASA/TM-2010-216089 4

0.84T^

1.01.4 0.781.82.22.6110 0.72

0.66

0.60

0.54

0.48

0.42int

0.36

40

3f

3(

02,1

m

m

a 2(d

1(

5 10 15 20 25 30 35

Corrected flowFigure 4.-One-spool core turboshaft engine compressor performance map and operating points.

0.9

0.8 _^-•

0.7

0.6.

-- - --- - - - --

w 0.5

1 -------

0.3 Design point n

0.2 - - - -1 2 3 4 5 6 7

Pressure ratio2.4 -

2.2

i - --

3: 10-0 1.8 (1 = 65 ::. 30m 50

a 1.6 ^y Map scalars, - 70o ! We = 0.0296 : 90U 1.4 Eff = 0.940 - 110

1.2 n Design point PR = 0.324N = 0.0175 • 130

1501.0

1 2 3 4 5 6 7Pressure ratio

Figure 5.-One-spool core turboshaft engine turbine performance maps and operating points

NASA/TM-2010-216089 5

U.0

0.8

0.7

7 0.6

0.5

0.4

w 0.3

0.2

0.1

001 2 3 4 5 6 7 8

Pressure ratio14

1210

11 :'" R = 74 - 3010 Map scalars,9-0 We = 0.070

C) Eff = 0.933'90

8 ; 110PR = 0.252 130

7 - ° • Design point N = 0.0259 1506

1 2 3 4 5 6 7 8Pressure ratio

Figure 6.-One-spool core turboshaft engine power turbine performance maps and operating points.

9

9

16

12

8

4

0

-4

-8

-12

-16

Weights Dimensions

Bare engine weight 895.7 Engine length 63.1Accessories weight 50.0 Engine pod C. G. 21.9Engine weight 945.7 Engine max diameter 24.0Inlet and nacelle weight 96.4 Nacelle max diameter 28.7Total engine pod weight 1042.1 Total engine pod length 69.3

-7.3-3.7 0.0 3.7 7.3 11.0 14.7 18.3 22.0 25.7 29.3 33.0 36.7 40.3 44.0 47.7 51.3 55.0 58.7 62.3 66.0 69.7

Figure 7.-One-spool core turboshaft engine WATE output.

NASA/TM-2010-216089 6

Two-Spool Core Engine

Noting the challenging compressor design for a one-spoolcore, a two-spool core (with a free power turbine on athird spool) was modeled. A block representation is given inFigure 8. Compression was split with roughly equal enthalpychange for the low and high pressure compressors, whichresulted in a low pressure compressor (LPC) pressure ratio of9.2 and a high pressure compressor (HPC) pressure ratio ofalmost 3.3 (maintaining the same engine overall pressure ratioof 30). Results are given in Table 3 detailing the flightcondition and pertinent operating characteristics for the lowpressure spool turbomachinery and power turbine. Thispreliminary analysis indicated that the two-spool coreturboshaft engine could achieve the horsepower requirementsfor all flight points (and power turbine mechanical rpms).Operation of the high pressure spool is essentially at constantcorrected rpm and correct mass flow rates (less than 2.3percent variation in the compressor and less than 1 percent inthe turbine aerodynamic values). The high pressure spool hada very minor physical over-speed (1 percent above the design

value) at hover, and was only operating at 90 to 97 percent ofthe physical design speed for cruise (the difference incorrected or aerodynamic versus physical rpm is from thereduction in component entrance temperatures going from hot-day hover to standard-day altitude conditions).

The operating points for the low pressure spoolturbomachinery and power turbine are shown in Figures 9 to11. As seen in Figure 9, the low pressure compressor correctedspeed varies a maximum of 7 percent from the designcondition; this maximum point is a result of the enginematching response to the hover condition on a hotter thanaverage day (45 °F hotter at 2;000 ft altitude). As shown inFigure 10, there is also essentially no variation in the lowpressure turbine corrected mass flow (<1 percent), althou gh itspressure ratio varies slightly (-4 to 6 percent from the designvalue). The operation of the power turbine is very similarbetween this two-spool core engine (Figure 11) and the one-spool core version (Figure 6). This was expected with thesimilarity in vehicle power requirements and gas properties(mass flow, temperature and pressure) provided by the core ofeach engine, therefore, the previous analysis still applies.

TABLE 3.-OPERATION OF TWO-SPOOL CORE TURBOSHAFT ENGINE OVERREFERENCE MISSION PROFILES AT VARIOUS POWER TURBINE RPMS

Altitude, Speed, Vtip, fps AT, Meet LP shaft rpm LPC LPT T4, PT HP,/ We-PTft knots (rotor) F lip? (achral/design) Nc Nc °F Nc Wturb^ (Ibnl/s)0 0 650 -27 Yes 1.02 99.2 100.1 3103 98.3 263.9 11.96

Mission 12.000 0 650 -45 Yes 0.97 93.6 99.5 2826 102.3 208.5 11.51

(100% PT 1prn) 28,000 303.4 350 0 Yes 0.89 96.7 99.9 2319 112.7' 201.8 11.95(75% PT rpm) Yes 0.89 96.9 100.6 2295 85.1' 200.2 12.18(50% PT rpm) Yes 0.91 98.4 101.4 2333 60.8 190.5 12.54

Mission-')5.000 0 650 -36 Yes 0.97 95.3 99.7 2817 102.7 218.5 11.66

(100% PT rprn) 28,000 330 350 0 Yes 0.91 98.0 100.1 2413 110.8' 219.5 12.06(75% PT rpm) Yes 0.91 98.5 100.8 2397 83.5' 216.3 12.32(50% PT rpm) Yes 0.93 100.3 101.6 2455 59.4 205.6 12.62

(50% PT rn) End 28,000 330 350 0 Yes 0.92 99.2 101.5 2397 60.1 199.1 12.60'Actual power turbine horsepower hnass flow (HP/lbm/sec)'Multispeed gearbox required to match rotor and power turbine rpnrs.

HPC bleeds for turbine cooling

Figure 8.-Block representation of a two-spool core turboshaft gas turbine model.

NASA/TM-2010-216089

12-

10 -I

0 8

mV

6a

4^

2 -I

0.84

1.01.41.8 0.782.22.6110 0.7205

0.66

0.60

0.54

0.48

In point^- 0.42

5 10 15 20 25 30 35Corrected flow

Figure 9.-Two-spool core turboshaft engine low-pressure compressor performance map andoperating points.

U. t)

0.80.7

0.6U

0.5

0.4w 0.3

0.2

0.1

001.0 1.5 2.0 2.5 3.0 3.5

Pressure ratio5.5

0 4.5 104.0 R = 74 30

0 3.5-y Map scalars, _ _ 70

o Wc = 0.0267 90v 3.0 / Eff = 0.923 a 110n Design point PR = 0.972 0 1302.5 N = 0.0194

1502.0

1.0 1.5 2.0 2.5 3.0 3.5Pressure ratio

Figure 10.-Two-spool core turboshaft engine low-pressure turbine performance maps and operating points

NASA/TM-2010-216089 8

n Design point

2 3 4 5 6 7 8

9 10Pressure ratio

/i

10

R = 74 ....... 3050Map scalars,70

We = 0.067~ 90Eff = 0.933' 110

PR =0.24 m 130 N = 0.026 150

n Design point

0.9

0.8

0.7

0.6

0.5

0.4

w 0.3

0.2

0.1

0.01

13

12

11

100W 9U

80U 7

6

5

....

a' --------

i

1 2 3 4 5 6 7 8 9

10Pressure ratio

Figure 11.—Two-spool core turboshaft engine power turbine performance maps and operating points.

Engine gas flow path dimensions and weights weregenerated assuming an all axial compression system; outputfrom the WATE analysis is shown in Figure 12. It resulted inan engine plus accessories total weight of 891 lb (about11 percent lighter than assumed in the LCTR2 study).Separating compression into two separate spools allowedsome reduction in compressor and turbine weight for the coreengine, although the core total shaft weight increased (the sumof the two, lighter core shafts required weighed 80 lb morethan the one core shaft of the one-spool concept). It alsoenabled a slightly larger compression exit blade height(0.65 in., almost 14 percent higher). Very small, latter stageblade heights are still an issue for this concept. This two-spoolcore concept also adds the complexity of a third shaft, whichmight be an issue considering the limitations on the third shaftdiameter size available imposed by the other two shafts andtorque. The third shaft might not be an issue for an aft turbinepower assembly, if such an option can be reasonablyintegrated with the drive system.

To get away from the small axial stages at compressionsystem exit, another engine model was investigated in which

the axial hi gh pressure compressor from the previous WATEiteration was replaced with a centrifugal stage. Althou gh thiscentrifugal stage is operating at higher temperatures thancentrifugal stages in present engines; it is worthwhile tounderstand centrifugal stages with increased temperaturecapability mi ght enable good solutions to the LCTR2 enginerequirements- As shown in Figure 13, it is very similar to thetwo-spool core, all axial, except for the hi gh pressurecompressor and combustor arrangement. Its weight at 1051 lbis a bit more than the other two engine concept weights,exceeding the LCTR2 wei ght requirement by 5 percent, andalleviates the small axial compressor blade height issue. Theadditional weight versus the two-spool, all axial engine isfrom the centrifugal compressor and combustor. Conservative,heavy materials were assumed, based on the high temperatureenvironment for these components. This additional weightcould be mitigated through better materials or careful engineflow path design, but those efforts are deferred to follow-onstudies. The engine layout also has the same potential issuewith its third shaft as previously discussed.

NASA/TM-2010-216089 9

-7.3-3.7 0.0 3.7 7.3 11.0 14.7 18.3 22.0 25.7 29.3 33.0 36.7 40.3 44.0 47.7 51.3 55.0 58.7 62.3 66.0 69.7 73.3

141210

86420

-2-4-6-8

-10-12-14

Weights Dimensions


141210

86420

-2-4-6-8

-10-12-14

-7.3-3.7 0.0 3.7 7.3 11.0 14.7 18.3 22.0 25.7 29.3 33.0 36.7 40.3 44.0 47.7 51.3 55.0 58.7 62.3 66.0 69.7 73.3

Figure 12.-AII axial, two-spool core turboshaft engine WATE output.

Weights Dimensions


Figure 13.-Axicentrifugal high-pressure compressor, two-spool core turboshaft engine WATE output.

NASA/TM-2010-216089 10

Conclusions

An engine thermodynamic, flow path and weight analysishas been performed to help define relevant areas of interest toachieve engine performance to meet LCTR2 performancegoals and enable a new vehicle class and operations into itsparticular niche in the national airspace. Studies suggest thatmodern ; large gas turbine engine temperatures and efficienciesare needed for the next generation of 7,500 hp class turboshaftengines. There has already been some recent work to look atdifferent compression configurations for this specific engine.Further efforts are needed to guide research to develop andverify designs for axial and centrifugal compressors to operateat temperatures, efficiencies, and sizes beyond the presentstate of the art. Turbine technology will also be needed tomaintain high efficiencies, reduced cooling requirements, andespecially verify efficient operation over a wide power turbinespeed range.

References1. Johnson, W., Yamauchi, G.K., and Watts, M.E., "NASA Heavy

Lift Rotorcraft Systems Investigation," NASA/TP-2005-213467, Sep. 200.5.

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4. Jones, Scott M., "An Introduction to ThermodynamicPerformance Analysis of Aircraft Gas Turbine Engine CyclesUsing the Numerical Propulsion System Simulation Code,"NASA/TM-2007-214690.

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01-02-2010 Technical Memorandum4. TITLE AND SUBTITLE 5a. CONTRACT NUMBERGas Turbine Characteristics for a Large Civil Tilt-Rotor (LCTR)

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6. AUTHOR(S) 5d. PROJECT NUMBERSnyder, Christopher, A.; Thunman, Douglas, R.

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14. ABSTRACTIn support of the Fundamental Aeronautics Program, Subsonic Rotary Wing Project ; an engine system study has been undertaken to helpdefine and understand some of the major gas turbine engine parameters required to meet performance and wei ght requirements as defined byearlier vehicle system studies. These previous vehicle studies will be reviewed to help define gas turbine performance goals. Assumptionsand analysis methods used will be described. Performance and weight estimates for a few conceptual gas turbine engines meeting theserequirements will be given and discussed. Estimated performance for these conceptual engines over a wide speed variation (down to 50percent power turbine r m at high for ue) will be presented. Finally, areas needin g ftuther effort will be suggested and discussed.15. SUBJECT TERMSTurboshafts; Gas turbine engine, Performance prediction, Weight

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